Low sonic boom inlet for supersonic aircraft

ABSTRACT

All-internal compression inlets for supersonic aircraft, with variable geometry systems and shock stability bleed systems provide high performance, large operability margins, i.e. terminal shock stability that reduces the probability of inlet unstart, and contribute little or nothing to the overall sonic boom signature of the aircraft. These inlets have very high internal area contraction or compression and very low external surface angles. All shocks from the internal inlet surfaces are captured and reflected inside the inlet duct, and all of the external nacelle surfaces are substantially aligned with the external airflow. The inlet shock stability system consists of bleed regions that duct bleed airflows to variable area exits with passive or active exit area controls. This reduces the risk of inlet unstarts by removing airflow through a large open throat bleed region to compensate for reductions in diffuser (engine) corrected airflow demand. Because the stability bleed is not removed until the inlet terminal shock moves upstream over the bleed region, the necessary normal shock operability margin is provided without compromising inlet performance (total pressure recovery, and distortion).

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of U.S. provisional application Ser.No. 60/235,359 filed on Sep. 26, 2000.

FIELD OF THE INVENTION

This invention relates to air intakes for flight vehicles and, moreparticularly, to air intakes for aircraft that are designed to fly atsupersonic speeds.

BACKGROUND OF THE INVENTION

Inlets for propulsion systems for high speed supersonic aircraft aredesigned to efficiently decelerate the approaching high-speed airflow tovelocities that are compatible with efficient airbreathing engineoperation and to provide optimum matching of inlet airflow supply toengine airflow requirements. Entrance airflow velocities to existingair-breathing engines must be subsonic; therefore, it is necessary todecelerate the airflow speed during supersonic flight. The airflowvelocities are slowed from supersonic speeds (above the speed of sound)to engine entrance Mach numbers that are subsonic (below the speed ofsound).

In aircraft propulsion systems having supersonic inlets, it is essentialthat the inlet decelerate the airflow in a manner that minimizes thepressure losses, cowl and additive drag, and flow distortion at theengine entrance. For supersonic inlets, efficient deceleration of thesupersonic velocities is accomplished by a series of weak shock wavesand/or isentropic compression, in which the speed is progressivelyslowed to an inlet throat Mach number of about 1.30. A terminal shockwave located near the inlet throat slows the airflow from supersonicspeeds (above the speed of sound) to subsonic speeds (below the speed ofsound). This terminal shock wave typically changes a Mach 1.3 flowcondition to a high subsonic flow level. Downstream of the terminalshock, the speed of the airflow is additionally slowed in the subsonicdiffuser of the inlet by a smooth transitioning of the flow duct from asmaller throat area to the larger area at the engine entrance.

Mixed-compression inlets, in which some of the supersonic compression ordeceleration in velocity is accomplished external to the duct and someof the compression is accomplished internally, have commonly beenproposed for supersonic aircraft that cruise at Mach numbers higher than2.0. Any inlet that accomplishes some of its compression internally issubject to an undesirable phenomenon known as inlet unstart. Inletunstart is characterized by an expulsion of the inlet terminal shockfrom the desirable location at the inlet throat station to a positionahead of the inlet cowling with an associated large increase in drag andlarge thrust loss. Unstart may also affect the aerodynamics of theaircraft.

Sonic boom is another factor that must be taken into account in thedesign of inlets of supersonic aircraft. Since economically viablesupersonic commercial aircraft must be able to operate supersonicallyover land, the inlet should contribute minimally to the sonic boomsignature of the aircraft. Therefore, the technical challenge for thedesigner of inlets for modern commercial aircraft is to provide a highperformance configuration that provides large operability margins(terminal shock stability to reduce the probability of inlet unstart),and to also identify a design that offers a reduction in the overallsonic boom signature of the aircraft. Mixed-compression inlets canefficiently decelerate the airflow while providing large operabilitymargins. However, the external compression, which is provided by acenterbody or cowl surface, radiates shock waves outward that contributeto the aircraft's sonic boom signature. These designs also have leadingedges that include an external surface at an angle to the local airflow.Oblique shock waves are generated by these surfaces, contributing to theaircraft's overall sonic boom problem. Over-land operation of commercialsupersonic aircraft requires that the sonic boom signature from theaircraft be reduced to acceptable levels. In order to achieve therequired acceptable boom levels, sonic boom contributions from eachcomponent on the aircraft must be reduced to the lowest possible level.

All-internal compression inlets are desirable from a sonic boomreduction standpoint, because they may be designed with no oblique shockwaves generated by an external compression system that would contributeto sonic boom signature. However, attempts to design these inlets havebeen generally unsuccessful, primarily due to large amounts of bleedrequired for inlet starting and started operation. Since these designstypically utilized fixed geometry, large amounts of bleed were necessaryto provide the effective flow area ratio from the inlet entrance toinlet throat to allow the inlet to start (establish a supersonic flowfield from the inlet entrance to the inlet throat). Large amounts ofbleed were also necessary during normal operation because these inletsdid not incorporated a stability system. This trend is typical of inletsthat do not incorporate a stability system. Adequate inlet stabilitymargins for inlet operation prior to unstart can only be provided by thefixed geometry bleed systems by prohibitively bleeding large amounts ofbleed airflow during normal operation. The development of a low sonicboom aircraft therefore requires an innovation in supersonic inletdesign.

SUMMARY OF THE INVENTION

The inlets disclosed and claimed herein provide high performance, largeoperability margins, i.e. terminal shock stability that reduces theprobability of inlet unstart, and contribute little or nothing to theoverall sonic boom signature of the aircraft. The characteristics ofthese inlets include very high internal area contraction or compressionand very low external surface angles. The design concept of thisinvention is a very high to all internal compression inlet, in which allshocks from the internal inlet surfaces are captured and reflectedinside the inlet duct (no compression system shocks radiated external tothe inlet duct). Additionally, they allow all of the external nacellesurfaces to be completely or very nearly aligned with the external flow(zero external surface angles). These low profile external surfaces donot produce a shock wave that contributes to the sonic boom problem. Inthis invention, an all-internal compression inlet is combined with ashock stability bleed system. The innovative application of a shockstability bleed system can prevent inlet unstarts caused by bothinternal and external flow disturbances, and provide large shockstability margins, thereby making the all internal-compression, or nearall-internal compression inlets feasible for application to supersoniccruise vehicles.

The inlet shock stability system consists of bleed regions that ductbleed airflows to variable area exits. The stability system incorporateseither passive or active exit area controls. This system prevents inletunstarts by removing airflow through a large open throat bleed region tocompensate for reductions in diffuser (engine) corrected airflow demand.Because the stability bleed is not removed until the inlet terminalshock moves upstream over the bleed region, the necessary normal shockoperability margin is provided without compromising inlet performance(total pressure recovery, and distortion) and without requiringprohibitive amounts of performance bleed during normal inlet operation.Research has demonstrated that the utilization of a variable bleed exiton a large open throat bleed region can provide very large inletstability margins for both internal and external airflow variations. Theappropriate placement of a stability bleed system in the throat of anall internal-compression inlet makes the design of such a configurationfeasible.

This all internal-compression inlet concept is designed to provide thehigh performance and reliability required for a highly efficientsupersonic aircraft and minimally contribute to sonic boom signature.The unique feature of the proposed design is the utilization of anall-internal compression scheme combined with a shock stability system.This type of inlet offers the opportunity to consider external surfacesthat are substantially aligned with the approaching airflow that willnot produce shock waves and the associated sonic boom. For inlets ofthis type, all of the supersonic compression is generated by thecontouring on the internal surface of the cowl since they do not employa centerbody.

Other features and advantages of this invention will be apparent tothose skilled in the art after reading the following detaileddescription and the accompanying drawings.

DRAWINGS

FIG. 1 shows an isometric view of a low sonic boom allinternal-compression inlet embodying this invention.

FIG. 2 presents a horizontal cross-sectional view of the inlet shown inFIG. 1, showing the internal cowl surfaces and an indication of theinlet aerodynamics.

FIG. 3 shows a downstream view of the inlet, i.e. in the direction ofairflow through the inlet, rotated 90° from FIG. 1 for ease ofcomparison with FIG. 2.

FIG. 4 presents a vertical cross-sectional view of the inlet that showsthe internal contours on the top and bottom surfaces of the inlet.

FIGS. 5 through 9A show cross-sectional views of the inlet.

FIGS. 10 through 10-D, 11 and 11-A present cross-sections of the inletthat show the cowl surfaces in the on-design (supersonic cruise)position and in the most off-design (low-speed) collapsed condition. Thedesign position is presented in FIG. 10 and the off-design position isshown in FIG. 11. Stability bleed regions are also depicted in FIG. 10.

FIGS. 10E through 10H are the same cross-sectional views of the inlet asshown in FIG. 10-10D, with the addition of an exit control valve.

FIGS. 12 through 14 show a mechanical mechanism to provide variablegeometry for a two dimensional (i.e. an inlet of rectangularcross-section in which the external surfaces from the leading edges tothe inlet throat are composed of flat or contoured plates) supersoniccruise inlet utilizing all internal-compression.

FIG. 15 shows an alternate leading edge for the top and bottom surfaces.

FIGS. 16 through 18 present approaches to adjust the top and bottomsidewalls of an inlet that is sized to meet the airflow demand of anengine with a requirement for a very low entrance Mach number.

FIG. 19 presents a configuration similar to the inlet of FIG. 1 with theleading edges of the cowl staggered.

FIG. 20 presents an alternate bifurcated inlet configuration thatutilizes the staggered concept of FIG. 19 in a back-to-back arrangement.

DETAILED DESCRIPTION

The basic inlet concept is presented in FIGS. 1 through 14. FIG. 1 showsan isometric view of the inlet, referred to generally as 1, and FIGS. 2through 4 present cross-sections of the configuration. The isometricsketch in FIG. 1 depicts a supersonic inlet 1 in which all the externalsurfaces are flow-aligned, i.e aligned with the airflow approaching theinlet. The airflow approaching the inlet is substantially parallel tothe inlet centerline; therefore, surfaces that are flow-aligned with thefreestream airflow are also parallel to the inlet centerline. Theinitial external cross-sectional shape of the inlet is rectangular andthen transitions as indicated by the surfaces 21 to a round nacelle atthe downstream end 10. If the propulsion system uses a square orrectangular nozzle, transitioning of the inlet surfaces, as shown bysurface 21 in FIG. 1, to a round nacelle is not required; therefore, therectangular cross-section would be continued to the end of the nacelle,station 10. This inlet 1 is composed of four surfaces: the sidewalls 55and 56 and top and bottom surfaces 53 and 52, respectively, of theinlet. As shown in FIG. 3 (rotated 90° relative to FIG. 1 for ease ofcomparison with FIG. 2), these surfaces (55, 56, 52 and 53) provide theinternal channel 51 to duct the captured airflow 77 through the inlet tothe exit station 10.

Referring to the horizontal cross-sectional view in FIG. 2, inlet 1 usesa low-angle (typically about 5° or less relative to the incomingairflow) initial compression wedge 3 on the internal cowl compressionsurface 2, which generates an initial oblique shock 4, i.e. a shock wavewith an angle less than 90° to the surface that is radiated out from theleading edge of from any compression surface angle. For example, a 5°wedge in a Mach 2.4 airstream generates an oblique shock wave with a28.73° angle to the incoming airflow. This internal cowl compressionsurface 2 includes the initial low angle wedge 3, an isentropic contour5, a throat section 6 (minimum cross-sectional area), and a subsonicdiffuser 7. Isentropic compression refers to a compression process thatis generated by a continuous curvature of the compression surface inwhich the airflow is progressively compressed or decelerated with noloss in the total pressure of the airstream. Isentropic compression canbe approximated by using a series of small angle changes to develop theoverall required compression. The isentropic compression contour 5provides the additional required supersonic compression or decelerationfrom the initial wedge 3 to the inlet throat section 6.

The isentropic compression flow field is depicted by the Mach waves 8.For example, in a typical supersonic transport installation, operatingat supersonic design conditions of about Mach 2.4, the supersonicairflow will have decelerated to about Mach 1.3 when it reaches thethroat 6. A normal (terminal) shock 9 at the inlet throat 6 willtypically further decelerate the airflow to about Mach 0.8. The subsonicairflow downstream of the terminal shock 9 continues to decelerate inthe subsonic diffuser 7 that extends from the inlet throat 6 to thediffuser exit station 10.

The internal inlet duct 51 is rectangular to a location just downstreamof the inlet throat 6 and then transitions to a circular cross-sectionat a station just upstream of the engine location 10. Tangent lines 11that are created by filleting the corners are shown. The subsonicdiffuser contains a break in the contour that provides an opening 12 toa typical overboard bypass system (not shown). As indicated in thedownstream view of the inlet presented in FIG. 3, the initial inletexternal surfaces are 16, 17, 18, and 19. FIG. 2 shows that externalsurfaces 16 and 17 are at 0° (flow aligned).

A downstream view of the inlet configuration is presented in FIG. 3. Thedistance between the internal surfaces 14 and 15 is equal to the enginediameter 61. These internal surfaces are also shown in FIG. 4. The topwall 53 is composed of an inner wall 14 and an exterior surface 18.Surface 14 exhibits an initial small compression surface angle to theincoming airflow 77 that is captured by the inlet. This small internalangle is necessary because the external angle for surface 18 is about0°. This small internal compression angle for surface 14 results in aweak shock wave 54. Proceeding downstream from the initial wedge,surface 14 then transitions to an axial direction with an expansion ofthe flow field. This expansion is represented by an initial expansionwave 64 and a final expansion wave 65. This internalcompression—expansion created by surface 14 and by the identicalopposite surface 15 should have very little effect on the overall inletcompression process. The airflow conditions approaching the inlet throatterminal shock 9 should mainly be the result of the compression systemcreated by the cowl surface 2 as shown in FIG. 2.

FIG. 5 shows the locations of several cross-sections (A—A to D—D) on theinlet 1. Cross-sectional views for these cross-sections are presented inFIGS. 6 through 9. Again as for FIG. 3, note that the cross-sections arerotated for ease of comparison with FIGS. 2 and 5. Cross-section A—A isshown in FIG. 6. In FIG. 6, both the internal duct (composed of surfaces2, 14 and 15) and the external shape (composed of surfaces 16, 19, 17,and 18) are rectangular. The shape is similar for FIG. 7 (cross-sectionB—B, FIG. 5) except the distance between the cowl surfaces 2 show therestriction of the duct area in the throat (minimum area) of the inlet.FIG. 8 shows the transitioning of the inlet to circular, both internallyand externally. The external surfaces are transitioned by the circulararcs 21, and the internal surfaces are transitioned by the circular arcs20. FIG. 9 shows a cross-section near the exit of the inlet in whichboth internal and external contours are circular. FIG. 9A shows across-section near the exit of the inlet in which both internal andexternal contours are elliptical.

This inlet utilizes a significant amount of isentropic compression. Thebenefits of isentropic compression and a throat Mach number of about 1.3will result in excellent total pressure recovery. In addition, theoverall reduction in performance due to boundary layer will be lower foran all-internal compression inlet than for of a conventionalmixed-compression inlet, since the basic inlet of this disclosure doesnot employ a centerbody. Inlets must provide a range of mass flows overwhich they can operate without the occurrence of an inlet unstart.Traditional performance boundary layer bleed systems can provide only asmall operability margin. Since this margin is generally not sufficient,additional margin is provided by operating at reduced performancelevels. A very high level of performance and an adequate operabilitymargin to prevent inlet unstart can be realized through the utilizationof a stability bleed system. This system allows operation of the inletat the optimum performance condition, and yet provides significant shockstability margins under conditions where an inlet unstart might tend tooccur, such as when the terminal shock moves upstream through the throatregion of the inlet due to a transient reduction in engine airflowdemand. The inlet stability bleed system compensates for changes indiffuser exit (engine) airflow demand by removing increasing amounts ofairflow from the inlet as the terminal shock moves upstream over theopen bleed regions that are located in the throat of the inlet. Thestability system functions to provide the necessary stability margins toprevent inlet unstart without prohibitive amounts of bleed during normalinlet operation by using variable area exit control valves that limitthe amount of bleed flow until increased bleed is required in responseto the upstream movement of terminal shock resulting from a transientdisturbance in inlet subsonic diffuser airflow.

An inlet throat stability bleed system is shown in FIGS. 10 through10-D. Uniformly distributed porous bleed is the preferred method toremove bleed airflow; however, any type of bleed opening can be used.For the preferred configuration, porous bleed surfaces are located inthe inlet throat section. Cowl bleed regions 23 are located in cowlsection 29, and sidewall bleed regions 24 are located in sidewalls 14and 15 (see FIGS. 10-B and 10-C). In the preferred embodiment, the openbleed regions 23 and 24 consist of the inlet surfaces with 0.125-inchholes drilled normal to the surface to obtain 40% open area (40%porosity). The bleed holes are located on 0.1875-inch centers with theholes in adjacent rows staggered to obtain a uniform distributedpattern. The preferred bleed surface would include a surface thicknessto hole diameter ratio of 1.0. The sidewall bleed 24 extends beyond thedesign cowl position so that bleed can be removed during off-designoperation. Folding compartment seals 44 are used to direct the inletbleed from the bleed surfaces (23 or 24) to exit passages andvariable-exit area controls, such as active or passive fast-actingvalves (not shown) at the bleed plenum exit, which control the amount ofbleed that is removed from the inlet.

FIGS. 10, 10-D, 11 and 11-A also illustrate one variable cowl geometrysystem that can provide the necessary variation of the internal surfacegeometry and well as changing the duct cross-sectional area at the inletthroat. Engine airflow demand varies as the flight vehicle speed changesfrom takeoff to supersonic cruise; therefore, a variation in the minimumduct area is necessary to accommodate the changes in airflow. Forefficient inlet operation, the internal surface geometry must also bechanged as the speed of the aircraft changes. This surface variation asthe flight vehicle speed changes allows the most optimum compression ofthe airflow that enters the inlet system. The internal inlet duct mustbe opened to a large area as illustrated in FIG. 11 during takeoff andfor low speed flight. As the flight vehicle accelerates to supersonicconditions, the variable geometry system is used to both provide theproper variation in inlet throat area as well as surface geometry.Comparison of the internal duct 51 geometry of FIGS. 11 and 10 shows thewide changes in the inlet geometry from takeoff to cruise speeds. Threehinge locations 25, 26, and 27 are shown in the Figures; however, thenumber of hinges may be any number suitable to provide proper cowlgeometry at off-design conditions. The variable cowl consists of anupstream section 28 hinged (25) at the upstream station and connected toadditional cowl sections 29 and 30 with hinges 26 and 27, with thedownstream end of the last section 30 including a guide pin 31 in agroove 32 (detail) to allow the length change for off-design operation,FIG. 10. The track 32 for the guide pin 31 is aligned to properlyposition the downstream end of the last cowl section 30. All cowlsections are hinged to the first cowl section 28. A sketch of the cowlin the off-design position is presented in FIG. 11. Note the change inposition of the downstream guide pin 31 between FIGS. 10 and 11.

Additional details of this variable cowl geometry scheme are presentedin FIGS. 12 through 14. Hydraulic actuators 43 are utilized to collapsethe cowl surfaces for off-design operation. These cylinders 43 arepinned 45 to bracket 33 that is attached to the outside surface 16 or 17at one end and pinned 46 at the other end to bracket 34 that is attachedto cowl surface 29. The hydraulic cylinders are attached to a commonfluid supply source so that uniform movement is obtained. Two actuatorsare shown in FIG. 14; however, any number could be used that would fitwithin the space available and effect the desired movement of the cowlsurfaces. While the hydraulic actuators provide the actuating power, theactual movement of the second cowl section 29 is controlled by ascissors arrangement that provides parallel positioning of the sectionfor any operating condition of the inlet. FIG. 12 shows that thisscissors arrangement is comprised of link bars 35 and 36 that are pinned37 and 38 to brackets 39 and 40 at the outer ends and pinned to frame 41at pin 42. Frame 41 is also shown in the isometric sketch of FIG. 14.The off-design position of the cowl 29 is shown in FIG. 13. As indicatedin a comparison of the cowl 29 vertical positions between FIGS. 12 and13, the inlet throat surface can be actuated to provide a significantincrease in duct area for off-design operation. The parallel throatsections 29 at design and off-design positions are shown in FIGS. 12 and13.

FIG. 1 shows an inlet with all external surfaces flow-aligned. However,this design requires the use of a small amount of compression on thewall of the inlet as shown in FIG. 4. Although small, as discussed forFIG. 4, this additional compression does result in some 3D flow in theinlet. The small internal compression wedges on the top and bottom inletwalls of the inlet generate a flow field that has a vertical crossflowcomponent. This crossflow component in the vertical plane of the inletinteracts with the crossflow component that is generated by thecompression surfaces in the horizontal plane. This interaction resultsin a 3D flowfield. This additional compression could be avoided if aconfiguration as presented in FIG. 15 is utilized. This design basicallyreverses the initial leading edge angle for the top and bottom wallsfrom a wedge angle on the inside surface to an angled wedge 22 on theexterior surface 71. Therefore, the resulting internal surface is flatwith no additional compression to create 3D flow effects. While thesmall angle on the exterior surface will generate a weak shock wave, itshould not significantly contribute to the sonic boom signature. Thus,the inlet configuration 81 of FIG. 15 offers the significant advantagesof the all-internal compression configuration with a small compromise inthe external surface sonic boom contribution for optimum internalaerodynamics.

The basic design problem of providing low external surface angles forlower supersonic cruise Mach number inlets is that the ratio of inletcapture area to engine face area gets smaller as the inlet design Machnumber decreases, particularly for inlets matched to jet engines thatrequire low entrance Mach numbers. For the Mach 2.4 inlet design that ispresented in FIGS. 1 through 14, sizing of the inlet capture area tosupply airflow to the jet engine 61 at an entrance Mach number of about0.4 provides an inlet 1 in which the angles of the external surfaces 16,17, 18 and 19 are 0° relative to the approach airflow 77 as shown inFIGS. 1 through 14. For this flow-aligned external-surface design, theexternal cross-sectional area of the inlet at the engine face station 10was increased by an amount necessary to provide a sufficient annularairflow passage between the outside of a jet engine 61 and the outernacelle 62 (FIG. 2) for cooling airflow around the exterior of theengine. The inlet of FIGS. 1 though 15 represent a design that has aminimum contribution to the sonic boom of a supersonic cruise aircraft.

If an engine is selected for a Mach 2.4 inlet that requires an entranceMach number less than about 0.4, all of the external surfaces cannot bealigned with the approach airflow. For low Mach number at the entranceto the engine, the engine area relative to the inlet entrance area willbe larger, and a slight external angle on the top 16 and bottom 17surfaces will result. Two transition schemes for the additional bulgeare shown in FIGS. 16 through 18. Since the largest cross-sectional areais at the inlet exit 10 (engine entrance), the largest bulge on theexternal surface will be at this location. To obtain a low boom designfor this inlet/engine combination, the external surface of the inlet istransitioned to the larger engine face area over a large distance on thenacelle upstream of the bulge, allowing a very small external angle andminimizing the resulting shock strength. The transitioning may have acircular arc shape 13 as shown in FIG. 16. As shown in FIGS. 17 and 18,the transitioning to the larger engine may extend along the entiresurface 72 as a curved flat surface 73 to the engine face 10. In FIG.18, only the surface contour 73 is shown. In either case, the low angledcontouring of the transitioning surface (13 or 73) would have little tono contribution to the sonic boom signature of the aircraft.

Several alternate configurations can be derived from the inlet designthat is shown in FIGS. 1 through 18 without departing from the basicdesign approach to identify a very low boom inlet configuration. Twosuch inlet configurations are shown in FIGS. 19 and 20. A staggeredinlet configuration 90 is presented in FIG. 19. Only the supersonicdiffuser of the inlet, from the leading edges 67 and 68 to the inletthroat station 97, is shown in FIG. 19. The subsonic diffuser for thisconfiguration would be similar to the one 7 shown in FIG. 2. This inlet90 is basically identical to the inlet 1 of FIG. 1 except the leadingedges 67 and 68 have been staggered to begin at different axialstations. This design offers the same performance and operability, wouldincorporate stability and variable geometry systems, and would have noexternal shock waves (no sonic boom) during operation at designconditions. Staggering of the leading edges offers some advantage forspilling airflow at off-design conditions. For the inlet 1 configurationof FIG. 2, in which the leading edges of the cowls 16, and 17 begin atthe same axial position, airflow cannot be spilled around the cowlingduring off-design flight speed conditions until the inlet unstarts. Uponunstart, airflow can spill around the cowling after it passes through astrong normal or bow shock that is located ahead of the inlet. Spillingairflow behind a strong normal shock has higher drag than supersonicspillage (spilling behind a supersonic oblique shock). Staggering ofcowl lips 67 and 68 of inlet 90 (FIG. 19) offers an unstarted inlet inwhich the normal shock is located ahead of lip 67 and an oblique shockis generated by lip 68. This oblique-normal shock combination offersmore efficient spillage of the airflow due to the reduction of thevelocity through the oblique shock prior to further deceleration throughthe normal shock.

An alternate inlet 50 developed by using the same design approach as forthe inlets 1 and 90 of FIGS. 2 and 19 is shown in FIG. 20. The inlet 50of FIG. 20 employs the staggered leading edge inlet design of FIG. 19 ina back-to-back arrangement to create a bifurcated configuration with avariable geometry centerbody and flow aligned external surfaces 76 and78. Inlet 50 of FIG. 20 is derived by placing surfaces 96 of two inlet90 from FIG. 19 together in such a way that a back-to-back bifurcatedinlet configuration is obtained. The internal duct rectangularcross-section at the throat of each of these inlets would betransitioned to a semi-circle at the exit 79 of the inlet to jointlyform a round entrance for a single engine. The large amount ofstaggering of the leading edges, leading edge 85 to 86 and leading edge85 to 87, for this configuration would provide nearly the sameoff-design spillage characteristics as a more conventionalmixed-compression inlet design. This inlet design 50 has all shock waves62 and 88 internal to the duct and all external surfaces 76 and 78 ofcowls 74 and 75 are flow aligned; therefore, this design, unlikeconventional designs, will not contribute to the sonic boom signature ofthe aircraft at design operating conditions.

The inlets defined in FIGS. 1 through 20 represent a new approach toinlet design. This invention offers inlet design options that can leadto new, more efficient, safer, and more environmentally friendlyaircraft. This inlet concept may offer integration options that were notpossible with more traditional inlets. This design approach can providean inlet configuration that will provide enabling technology for a quiet(low sonic boom), efficient, supersonic cruise aircraft.

While 2-dimensional inlet configurations have been described in FIGS. 1through 20, it will be evident to those skilled in the art that theconcept may be extended to the design of axisymmetric inlets withsimilar attributes and benefits.

It is understood that the invention is not limited to the specificembodiments herein illustrated and described, but may be used in otherways without departing from its spirit. Other embodiments of theinternal compression inlet described herein that suggest themselves tothose skilled in the art are intended to be covered by the claims ofthis disclosure which are as follows:

1. A supersonic inlet for use with an aircraft, said supersonic inletcomprising: an internal duct having an opening for receiving airflow anda throat section, said inlet further comprising one or more externalsurfaces wherein said one or more entire external surfaces are alignedwith the flow of the aircraft, said opening of said internal ductfurther comprised of a first and second leading edge, wherein saidleading edges are staggered in location; and said throat section of theinlet further incorporating a shock stability bleed system, wherein aportion of said airflow is removed from said internal duct so that ashock wave is maintained within said throat section.
 2. A supersonicinlet for use with an aircraft, said supersonic inlet comprising: aninternal duct having an opening for receiving airflow and a throatsection, said inlet having one or more external surfaces aligned withthe airflow of said aircraft, a centerbody positioned within said duct,and having a leading edge, the external surfaces being aligned with theairflow of the aircraft from the leading edge to an associated engine;said opening of said internal duct further comprised of a leading edge,wherein said leading edge of said duct is staggered in location withrespect to the leading edge of the centerbody; and said throat sectionof the inlet further includes a shock stability bleed system, wherein aportion of said airflow is removed from said internal duct so that ashock wave i maintained within said throat section.
 3. The inlet ofclaim 1 wherein all of the external surfaces are aligned with theairflow of the aircraft.
 4. The inlet of claim 1 wherein one or more ofthe external surfaces are aligned with the airflow of the aircraft atall operating conditions.
 5. The inlet of claim 1 wherein said internalduct comprises one or more movable compression surfaces.
 6. The inlet ofclaim 1 wherein said one or more external compression surfaces are fixedin location.
 7. The inlet of claim 2 wherein all of the externalsurfaces are aligned with the airflow of the aircraft.
 8. The inlet ofclaim 2 wherein one or more of the external surfaces are aligned withthe airflow of the aircraft at all operating conditions.
 9. The inlet ofclaim 2 wherein said internal duct comprises one or more movablecompression surfaces.
 10. The inlet of claim 2 wherein said one or moreexternal compression surfaces are fixed in location.
 11. A low boomsupersonic inlet for use with an aircraft engine, the supersonic inletcomprising: one or more internal surfaces forming an internal duct toprovide airflow to the aircraft engine and having an opening forreceiving airflow; and one or more external surfaces that are entirelyaligned with the flow of air to the inlet whereby the inlet externalshock waves that contribute to the sonic boom signature of the aircrafthave been substantially reduced.
 12. The low sonic boom supersonic inletof claim 11 wherein said internal surfaces provide supersoniccompression and are movable into two or more positions.
 13. The lowsonic boom supersonic inlet of claim 11 wherein said internal surfacesextend from the leading edge to the engine entrance, said internalsurfaces comprising one or more compression surfaces for providingsupersonic compression of the air to the inlet throat and subsonicdiffusion of the air to the engine entrance.
 14. The low sonic boomsupersonic inlet of claim 11 wherein said internal duct furthercomprises a throat section having a shock stability bleed systemcomprised of one or more bleed passageways wherein a portion of saidairflow is removed from said internal duct through said one or morebleed passageways so that an airflow shock wave is maintained withinsaid throat region.
 15. The low sonic boom supersonic inlet of claim 11wherein all of the compression surfaces are located within said internalduct.
 16. The low sonic boom supersonic inlet of claim 15 wherein excessair not required by the engine is ducted out of the inlet through anoverboard bypass system.
 17. The low sonic boom supersonic inlet ofclaim 11 wherein said internal duct has a rectangular cross-sectionalshape.
 18. The low sonic boom supersonic inlet of claim 11 wherein saidinternal duct has a elliptical cross-sectional shape.
 19. The low sonicboom supersonic inlet of claim 1 wherein said internal duct has arectangular cross-sectional shape.
 20. The low sonic boom supersonicinlet of claim 1 wherein said internal duct has a ellipticalcross-sectional shape.
 21. The low sonic boom supersonic inlet of claim2 wherein said internal duct has a rectangular cross-sectional shape.22. The low sonic boom supersonic inlet of claim 2 wherein said internalduct has a elliptical cross-sectional shape.
 23. The low sonic boomsupersonic inlet of claim 1 wherein said shock stability bleed systemfurther comprises bleed passages having a variable area exit.
 24. Thelow sonic boom supersonic inlet of claim 1 wherein said throat sectionof said inlet further comprises one or more movable internal sidewallsin the throat section for varying the throat area.
 25. The low sonicboom supersonic inlet of claim 1 wherein the interior surfaces of theinternal duct have continuous surfaces from the opening to the exit ofthe inlet.